Interferometer type homing head for guided missiles

ABSTRACT

Two detecting systems are used to generate yaw and pitch error signals for obtaining a collision course with a target, and the systems are interconnected to exchange information.

United States Patent [1 1 Schaefer INTERFEROMETER TYPE HOMING HEAD FORGUIDED MISSILES V Inventor: Jacob W. Schaefer, Watchung, NJ.

Assignee: The United States of America as represented by the Secretaryof the Army Filed: Nov. 23, 1966 Appl. No.: 596,736

References Cited UNITED STATES PATENTS 8/l965 Kent et al. 343/117 A X3/1966 Simon et al 244/3.l9

[ June 19, 1973 OTHER PUBLICATIONS Sommer, Howard H., I.R.E.Transactions on Aeronautical and Navigational Electronics, Vol. ANE3,N0. 2, June, 1956, pp. 67-70.

Primary Examiner-Benjamin A. Borchelt Assistant ExaminerRichard E.Berger AttorneyHarry M. Saragovitz, Edward J. Kelly, Herbert Ber] andRobert C. Sims 57 ABSTRACT Two detecting systems are used to generateyaw and pitch error signals for obtaining a collision course with atarget, and the systems are interconnected to exchange information.

6 Claims, 3 Drawing Figures I 3l CONVERTER TAN P MOTOR c 1 l l 20 l r I30 34 CONVERTER SEC P SERVO SEC Y 36 Y ERRoR ERRoR P SENSOR SENSOR 35GUIDANCE PAIENIEH 3.740.002

sum-1 N 2 TARGET PITCH' PLANE YAW PLANE \SSQF'. RV FIG. I

I 3l CONVERTER TAN? MOTOR 0 l l 20 v I l /30 34 CONVERTER SEC P SERVOsec Y I l L. ERROR ERROR p 35 SENSOR SENSOR GUIDANCE Jacob W. Schoefer,INVENTOR.

BY M .1. MW

PATENIEB 9 SHEEIEN2 PITCH SOLUTION INVENTOR.

B Elwin J. /W

Jacob W. Schaefer,

YAW SOLUTION 9 N m 4 mm MT AY m% RG A KHAN.H R 5 YT m 3 MM M 00 VI T .YH m .H m m A 3 S F W G F 5 4 7 4 L0 AV NR E 5 W 0 Y R0 ET 5 GUIDANCEINTERFEROMETER TYPE HOMING HEAD FOR GUIDED MISSILES The use of a hominghead will increase the effectiveness of most of the intercept missiles.A homing head will also incrase a systems effectiveness againstformations of targets, and it will increase the probability of killagainst single targets at very long ranges. However, there is often alimited amount of space available for a homing head. Therefore, there isneed for a homer which has a very simple antenna system so as to permitit to be placed in the nose compartment of a small missile. The systemwill have to be able to detect and track targets up to 45 off the noseof the missile.

It is an object of the invention to provide an interferometer typehoming head for use in missiles.

It is a further object of the present invention to provide homing headwhich has a simple compact antenna system.

A still further object of this invention is the provision of a homingsystem which uses proportional navigation with respect to its own axis.

The invention further resides in and is characterized by various novelfeatures of construction, combinations, and arrangements of parts whichare pointed out with particularity in the claims annexed to and forminga part of this specification. (Iomplete understanding of the inventionand an introduction to other objects and features not specificallymentioned will be apparent to those skilled in the art to which itpertains when reference is made to the following detailed description ofa specific embodiment thereof and read in conjunction with the appendeddrawing. The drawing, which forms a part of the specification, presentsthe same reference characters to represent corresponding and like partsthroughout the drawing, and wherein:

FIG. 1 illustrates the geometry for a homing solution of the presentinvention;

FIG. 2 illustrates in block diagram the overall system of a preferredform of the invention; and

' FIG. 3 illustrates in greater detail the yaw solution structure of theinvention.

Nearly all homing heads operate on the principle of proportionalnavigation. This term is used to describe a system that measures thedirection of the target, in earths coordinates, and steers the missilein a manner to keep that direction constant. If this criteria isrealized in both steering coordinates the missile is on a collisioncourse, and the direction of the tqrget remains unchanged untilintercept. One difficulty arises from the fact that the direction of thetarget must be measured irrespective of missile attitude changes. Thiscan be accomplished by placing the homing head on a stabilized platform.Real target motion can also be obtained by subtracting missile bodymotions from the observed target motion. An interferometer type homerutilizes the latter technique. However, the interferometer will employan approximate manner of subtracting missile motion from the observedtarget position. This is a cause of considerable concern as it is asource of instability in the steering loop. An approximate solutionwould be especially detrimental in a system where the angle of attackper unit of acceleration is relatively large. This would mean thatrelatively large missile motions would have to be subtracted and thecorrections would have to be much more exact if instability is to beavoided. The homing device of the present invention provides a completesolution for the missile motion problem and will eliminate it as a causeof instability. FIG. 1 is a diagram that illustrates the geometry of theproblem. A set of rectangular coordinate axes is illustrated, one ofwhich is along the center line of the missile structure. The other twoaxes are determined by the location of the antenna pairs. These causethe other two axes to be in the yaw and pitch planes. The interferometertype device of this invention will measure the angles in the slantplanes containing the target and a pair of the antennas of theinterferometer. In the case of the solution for the yaw plane, the angleis Y. However, the guidance of the missile is steered in the planecontaining the angle Y. The same is true for the slant angle P and thepitch angle P. The relationships of these angles is shown in thefollowing equations:

tan Yltan Y= cos P tan Y= sec P tan Y tan P'ltan P cos I tan P sec Y tanP FIG. 2 shows a block diagram of the equipment necessary to obtain Yand P and solve the equations to produce error indications to guidance5. Two pair of antennas 7 and 8, and 9 and 10 are used to determine theposition of the target. The pair 7 and 8 are aligned in the yaw planewhile the pair 9 and 10 are aligned in the pitch plane. Three antennasmay be used where one of the antennas (for example 7) will be used inboth the yaw and the pitch plane. Phase detectors 11 and 12 are providedat the junctions of the pairs of wave guides 14 and 15, and 16 and 17.The outputs of the phase detectors are used to drive servo motors l8 and19 by way of amplifiers 20 and 21. The motors 18 and 19 in turn drivephase shifters 23 and 24 so as to null the phase difference at thedetectors. The shaft 26 position of motor 18 is proportional to sin Y,and the shaft 27 position of motor 19 is proprotional to sin P. Thesepositions are converted by converter means including converters 30 and31 to outputs which represent tan Y and tan P. However, in order forconverters 30 and 31 to properly operate they must have as inputs thesec P and the sec Y respectively. This is obtained by servo units 33 and34 in the converter means. The servo units which convert their inputs tosec Y and sec P respectively. Servos 33 and 34 also have outputs of Yand P which are fed through error sensors 35 and 36 to the guidance 5for steering of the missile. Error sensors 35 and 36 may each be made upof differentiator 45, amplifier 47, and rate gyro 49 connected as shownin FIG. 3.

FIG. 3 shows the solution of the yaw equations in greater detail. Thedetail of the pitch equipment is identical to that of the yaw equipment,except for being different in the phase placement. As can be seen fromFIG. 3 the received radiation from a target will arrive at differenttimes on antennas 7 and 8. This difference in time is proportional to Yand can be determined by the difference in the phase of the signalsreceived by the antennas. A phase detector 11 provides an error outputsignal to amplifier 20 which drives a servo motor 18 which in turnpositions a phase shifter in the wave guide 15 of antenna 8. This loopwill find a null and the position 4: of the shaft 26 of motor 18, and 4)will be a measure of the relative phase difference of the receivedsignals on antennas 7 and 8. The relationship of the shaft position (1;and Y is (b 211- L/k) sin Y where d) shaft position A wave length ofradiation L distance between antennas.

Equation (5) must be manipulated to extract tan Y for use in equation(2).

sin Y'=)t/21rL Y sin (k l2 7: L)

tan Y sin ()t /2 w L) substituting in equation (2),

tan Y= sec P [tan sin (A (11/2 1r L)] From equation (6) it can be seenthat a voltage proportional to tan Y can be developed by a potentiometerif the shaft rotation is d), the card shape is the function of tan sin(A l2 n L), and the card is supplied by a voltage proportional to sec P.This situation is provided in FIG. 3 by the potentiometer 40 which hasits shaft rotated by the motor 18, has a card shape 43 in accordance tothe above function, and is supplied by a voltage proportional to sec Pfrom the pitch solution. Therefore, the output of potentiometer 40 isproportional to tan Y. This is fed to a computing type of servo 33 whichhas outputs of Y and sec Y. This computing type of servo could be of themeter movement type.

The output sec Y of servo 33 is sent to supply the potentiometer in thepitch solution. The output Y of servo 33 is sent through adifferentiator 45 so as to get the change in yaw direction of target Y.This is sent through amplifier 47 where the change in the missiles yaw His subtracted therefrom. The output of the amplifier is sent to the yawservo control for providing any needed correction to the missiles yawdirection. Rate gyro 49 provides the signal for the change in themissiles yaw. This will be an absolute value as it is only related tothe missile itself and not to the target or ground.

When the same is done for the pitch of the missile, then the guidance ofthe missile will keep the yaw and pitch directions of the targetconstant. This will cause the missile and the target to be on collisioncourses. The missile could contain a warhead which could be explodedupon command, contact, proximity, etc. The target could be illuminatedfrom the ground, by a transmitter within the missile, by the target'sjamming system, etc.

Variations in frequency of the secondary radiation from the target willcause errors in the target angular rate solution due to the change inthe wave length A in the equations. Errors of this sort will not effectthe accuracy of indicating when a true collision course has beenattained. However, they will affect over-all loop stability by degradingthe compensation for transient effects of body motion. Fortunately, thisaffect can be nullified automatically by the phase shifters. If one ofthe more common types of phase shifter is used, then it will operate ona delay principle instead of causing a specific angular effect. Theresulting angular dealy, therefore, is proportional to frequency, andwill tend to automatically compensate for any change in wave length.

A preferred embodiment of the invention has been chosen for purposes ofillustration and description. The preferred embodiment illustrated isnot intended to be exhaustive nor to limit the invention to the preciseform disclosed. It is chosen and described in order to best explain theprinciples of the invention and their application in practical use tothereby enable others skilled in the art to best utilize the inventionin various embodiments and modifications as are best adapted to theparticular use contemplated. It will be apparent to those skilled in theart that changes may be made in the form of the apparatus disclosedwithout departing from the spirit of the invention as set forth in thedisclosure, and that in some cases certain features of the invention maysometimes be used to advantage without a corresponding use of otherfeatures. It is, therefore, to be understood that within the scope ofthe appended claims, the invention may be practiced otherwise than asspecifically described. Accordingly, it is desired that the scope of theinvention be limited only by the appended claims.

I claim:

1. In a carrier means, a guidance system comprising a guidance means,first and second detecting means, each of the detecting systems havingfirst and second radiation detectors, respectively having inputs fordetecting a radiation and outputs which are measures of the radiation asreceived at their inputs, phase detecting means respectively connectedto the outputs of said radiation detectors, converter means connected tooutputs of each phase detecting means for converting phase differencesbetween the outputs of the respective first and second detectors intomeasures of the directions of the detected radiation relative to theradiation detectors, sensing means connected to outputs of saidconverter means for sensing changes of the directions of the radiationsrelative to the radiation detectors, the converter means of each of thedetecting systems being interconnected with each other to exchangeinformation, said carrier means containing said guidance means, and saidguidance means being connected to outputs of said sensing means forguiding said carrier means in a direction such that no change will besensed by said sensing means.

2. A guidance system as in claim 1 wherein adjustable phase shifters areconnected between corresponding radiation detectors and said phasedetectors, and positioning means are respectively connected to saidphase detectors and disposed to operate said phase shifters andconverters to provide similar phase relations between the outputs of therspective detecting systems.

3. A guidance system as set forth in claim 1, wherein said carrier meansis a missile, said first detecting system having its detectors alignedto respond to changes in the yaw direction of the radiation, and saidsecond detecting system having its detectors aligned to re spond tochanges in the pitch direction of the radiation.

4. A guidance system as set forth in claim 1, wherein each of saidconverter means comprises a potentiometer means having a movable armpositioned by a positioning means and a card which has a predeterminedspond to changes in the pitch direction of the radiation.

6. A guidance system as set forth in claim 5, wherein said sensing meanseach contain a differentiator connected to the output of the servoconverter, a rate gyro, and a combining means having inputs connected tooutputs of the differentiator and the rate gyro; and said combiningmeans having an output connected to said guidance means.

1. In a carrier means, a guidance system comprising a guidance means,first and second detecting means, each of the detecting systems havingfirst and second radiation detectors, respectively having inputs fordetecting a radiation and outputs which are measures of the radiation asreceived at their inputs, phase detecting means respectively connectedto the outputs of said radiation detectors, converter means connected tooutputs of each phase detecting means for converting phase differencesbetween the outputs of the respective first and second detectors intomeasures of the directions of the detected radiation relative to theradiation detectors, sensing means connected to outputs of saidconverter means for sensing changes of the directions of the radiationsrelative to the radiation detectors, the converter means of each of thedetecting systems being interconnected with each other to exchangeinformation, said carrier means containing said guidance means, and saidguidance means being connected to outputs of said sensing means forguiding said carrier means in a direction such that no change will besensed by said sensing means.
 2. A guidance system as in claim 1 whereinadjustable phase shifters are connected between corresponding radiationdetectors and said phase detectors, and positioning means arerespectively connected to said phase detectors and disposed to operatesaid phase shifters and converters to provide similar phase relationsbetween the outputs of the rspective detecting systems.
 3. A guidancesystem as set forth in claim 1, wherein said carrier means is a missile,said first detecting system having its detectors aligned to respond tochanges in the yaw direction of the radiation, and said second detectingsystem having its detectors aligned to respond to changes in the pitchdirection of the radiation.
 4. A guidance system as set forth in claim1, wherein each of said converter means comprises a potentiometer meanshaving a movable arm positioned by a positioning means and a card whichhas a predetermined function, connections from said movable arm to aservo converter, each servo converter having two outputs, one connectedto an input of the card of the potentiometer of the other detectingsystem, and the other output connected to the sensing means.
 5. Aguidance system as set forth in claim 4, wherein said carrier means is amissile, said first detecting system having its detectors aligned torespond to changes in the yaw direction of the radiation, and saidsecond detecting system having its detectors aligned to respond tochanges in the pitch direction of the radiation.
 6. A guidance system asset forth in claim 5, wherein said sensing means each contain adifferentiator connected to the output of the servo converter, a rategyro, and a combining means having inputs connected to outputs of thedifferentiator and the rate gyro; and said combining means having anoutput connected to said guidance means.